Modal noise reduction for gas turbine engine

ABSTRACT

An exemplary section of a gas turbine engine according to this disclosure includes, among other things, a first array of airfoils including a first number of airfoils, and a second array of airfoils downstream of the first array of airfoils. The second array includes a second number of airfoils. The second number of airfoils is at least 1.19 times the first number of airfoils thereby providing a predetermined modal.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/680,409 filed Apr. 7, 2015.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. The compressorsection compresses air and delivers it into a combustion chamber. Thecompressed air is mixed with fuel and combusted in the combustionsection. Products of this combustion pass downstream over turbinerotors. The compressor is typically provided with rotating blades, andstator vanes adjacent to the blades. The stator vanes control the flowof the air to blades. The arrangement between the stator vanes and theblades has an influence on the amount of noise (e.g., sound) generatedby the engine.

SUMMARY

An exemplary section of a gas turbine engine according to thisdisclosure includes, among other things, a first array of airfoilsincluding a first number of airfoils, and a second array of airfoilsdownstream of the first array of airfoils. The second array includes asecond number of airfoils. The second number of airfoils is at least1.19 times the first number of airfoils thereby providing apredetermined modal response.

In a further non-limiting embodiment of the foregoing section, thesecond number of airfoils is within a range between 1.19 and 1.55 timesthe first number of airfoils.

In a further non-limiting embodiment of the foregoing section, the rangeis defined by the following equation

${1 + \frac{M_{tip}\sin \; \theta}{1 - M}};$

where M_(tip) is a tip rotational Mach number, M is the Mach number intothe second array from a frame of reference of the second array, and θ isa stagger angle of the second array.

In a further non-limiting embodiment of the foregoing section, thestagger angle is the incline of a chord between a leading edge and atrailing edge of an airfoil relative to a direction parallel to anengine central longitudinal axis.

In a further non-limiting embodiment of the foregoing section, the rangeis further defined by:

$\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$

when the upstream array is a rotor and the downstream array is a stator;

$\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$

when the upstream array is a stator and the downstream array is a rotor;

where n is the harmonic of blade passing frequency, and k is theharmonic of vane passing frequency.

In a further non-limiting embodiment of the foregoing section, n and kare equal to 1.

In a further non-limiting embodiment of the foregoing section, thesection includes a plurality of arrays of airfoils, and a number ofairfoils in each array is at least 1.19 times a number of airfoils in animmediately upstream array.

In a further non-limiting embodiment of the foregoing section, thenumber of airfoils in each array is within a range between 1.19 and 1.55times the number of airfoils in the immediately upstream array.

In a further non-limiting embodiment of the foregoing section, the firstarray of airfoils is an array of stator vanes, and wherein the secondarray of airfoils is an array of rotor blades.

In a further non-limiting embodiment of the foregoing section, thesection is a low pressure compressor.

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes, among other things, a low pressure compressorincluding a first array of airfoils and a second array of airfoilsdownstream of the first array. The first array includes a first numberof airfoils and the second array includes a second number of airfoils.The second number of airfoils is at least 1.19 times the first number ofairfoils thereby providing a predetermined modal.

In a further non-limiting embodiment of the foregoing engine, the secondnumber of airfoils is within a range between 1.19 and 1.55 times thefirst number of airfoils.

In a further non-limiting embodiment of the foregoing engine, the rangeis defined by the following equation

${1 + \frac{M_{tip}\sin \; \theta}{1 - M}};$

where M_(tip) is a tip rotational Mach number, M is the Mach number intothe second array from a frame of reference of the second array, and θ isa stagger angle of the second array.

In a further non-limiting embodiment of the foregoing engine, thestagger angle is the incline of a chord between a leading edge and atrailing edge of an airfoil relative to a direction parallel to anengine central longitudinal axis.

In a further non-limiting embodiment of the foregoing engine, the rangeis further defined by:

$\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$

when the upstream array is a rotor and the downstream array is a stator;

$\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$

when the upstream array is a stator and the downstream array is a rotor;

where n is the harmonic of blade passing frequency, and k is theharmonic of vane passing frequency.

In a further non-limiting embodiment of the foregoing engine, n and kare equal to 1.

A method according to an exemplary aspect of this disclosure includes,among other things, controlling an interaction between a first array ofairfoils and a second arrays of airfoils during operation of a gasturbine engine by providing a number of airfoils in the second arraythat is at least 1.19 times the number of airfoils in the first array.The second array of airfoils is downstream of the first array ofairfoils.

In a further non-limiting embodiment of the foregoing method, the numberof airfoils in the second array is within a range between 1.19 and 1.55times the number of airfoils in the first array.

In a further non-limiting embodiment of the foregoing method, the rangeis defined by the following equation

${1 + \frac{M_{tip}\sin \; \theta}{1 - M}};$

where M_(tip) is a tip rotational Mach number, M is the Mach number intothe second array from a frame of reference of the second array, and θ isa stagger angle of the second array.

In a further non-limiting embodiment of the foregoing method, the rangeis further defined by:

$\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$

when the upstream array is a rotor and the downstream array is a stator;

$\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$

when the upstream array is a stator and the downstream array is a rotor;

where n is the harmonic of blade passing frequency, and k is theharmonic of vane passing frequency.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings can be briefly described as follows:

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a highly schematic view of a section of the gas turbineengine, and in particular illustrates a modal, cuton interaction betweenadjacent airfoil arrays.

FIG. 3 is another highly schematic view of the section of the gasturbine engine, and in particular illustrates a non-modal, cutoninteraction between adjacent airfoil arrays.

FIG. 4 graphically illustrates the relationship between sound powerlevel and the ratio between the number of airfoils in a downstream arrayversus the number of airfoils in a upstream array.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

FIG. 2 is a highly schematic view of a portion of a section of the gasturbine engine 20. In this example, the section is the low pressurecompressor 44. However, it should be understood that this disclosure maybe useful in other sections of the gas turbine engine 20.

In this example, the low pressure compressor 44 includes a plurality ofcompression stages, each including an array of rotor blades and an arrayof stator vanes. FIG. 2 shows two adjacent airfoil arrays. A first array60 includes a first number airfoils 62, which in this example may bestator vanes. A second array 64 of airfoils, which is immediatelydownstream of the first array 60, includes a second number of airfoils66. Continuing with the example, the airfoils 66 in the second array 64are rotor blades configured to rotate about the engine centrallongitudinal axis A.

The relationship between the number of airfoils in the first array 60and the number of airfoils in the second array 62 can be controlled toreduce engine noise. To this end, a concept known as “cutoff” has beenused in the design of compressors. “Cutoff” designs are typically usedin larger engines. In a cutoff configuration, the vane-blade ratio isset such that the blade passing frequency decays in the duct. Anotherknown concept is “high frequency.” This solution is typically beneficialin engines with relatively high speed low pressure compressors. Anotherconcept separate from “cutoff” and “high frequency” is called “cuton”(sometimes spelled “cut-on”). In a cuton configuration, the vane-bladeratio is set such that blade passing frequency propagates in the duct.

In this disclosure, the number of airfoils between adjacent arrays isselected such that, while not “cutoff,” there is a controlled “modal”interaction between the adjacent arrays. As will be appreciated from thebelow, the nature of the “modal” interaction reduces sound effectivelyeven though “cutoff” cannot be achieved.

In one example of this disclosure, each successive downstream array ofairfoils in the low pressure compressor 44 includes more airfoils thanthe upstream array. In particular, the number of airfoils in thesubsequent, downstream array is a factor of the number of airfoils inthe immediately upstream array. In one example, the factor F is the samefor each successive array, and is defined by the following equation:

$\begin{matrix}{F = {1 + \frac{M_{tip}\sin \; \theta}{1 - M}}} & ( {{Equation}\mspace{14mu} 1} )\end{matrix}$

where, for the aircraft's approach power conditions, M_(tip) is a tiprotational Mach number, M is the Mach number into the downstream arrayfrom the downstream array's frame of reference, and θ is a stagger angleof the downstream array. The factor F is a vane/blade ratio (V/B asdefined below) when the upstream array is a rotor and the downstreamarray is a stator. The factor F is a blade/vane ratio (B/V as definedbelow) when the upstream array is a stator and the downstream array is arotor. F can be defined in general by:

$\begin{matrix}{{{F = \frac{k\; V}{nB}};{{for}\mspace{14mu} {upstream}\mspace{14mu} {rotor}}},{{downstream}\mspace{14mu} {stator}}} & ( {{Equation}\mspace{14mu} 2} ) \\{{{F = \frac{kB}{nV}};{{for}\mspace{14mu} {upstream}\mspace{14mu} {stator}}},{{downstream}\mspace{14mu} {rotor}}} & ( {{Equation}\mspace{14mu} 3} )\end{matrix}$

where n is the harmonic of blade passing frequency, B is the number ofrotor blades in the blade row interaction, k is the harmonic of vanepassing frequency, and V is the number of stator vanes in the stator rowinteraction. Further, in one example, the stagger angle θ is the inclineof a chord 68 between a leading edge 70 and a trailing edge 72 of anairfoil relative to a line 74 parallel to the engine centrallongitudinal axis A. Alternatively, the stagger angle θ could be theangle of a line tangent to the leading edge.

In this example, the value for n is 1 because the first harmonic ofblade passing frequency has the most significant impact on noisereduction. The value for k is also 1, because this is the only cutonmode for the first harmonic of blade passing frequency.

The result of the above Equations 1, 2, and 3 provides maximum noisereduction when in a “cuton” state when the factor F is at least 1.19. Inparticular, the maximum noise reduction is achieved when the factor F iswithin a range between 1.19 and 1.55. That is, for a rotor blade passingfrequency, the number of airfoils in an array is within a range between1.19 and 1.55 times the number of airfoils in the immediately upstreamarray.

The range of the factor F is further defined by the following equations,which are the result of combining Equations 1, 2, and 3.

$\begin{matrix}{{{\frac{V}{B} = {\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )}};}{{{for}\mspace{14mu} {an}\mspace{14mu} {upstream}\mspace{14mu} {rotor}},{{downstream}\mspace{14mu} {stator}}}} & ( {{Equation}\mspace{14mu} 4} ) \\{{{\frac{B}{V} = {\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )}};}{{{for}\mspace{14mu} {an}\mspace{14mu} {upstream}\mspace{14mu} {stator}},{{downstream}\mspace{14mu} {rotor}}}} & ( {{Equation}\mspace{14mu} 5} )\end{matrix}$

In one example, the first array 60 includes 30 stator vanescircumferentially spaced around a disk. In that example, the secondarray 64 could include between 36 and 47 rotor blades. Continuing withthat example, a third array (not pictured) immediately downstream of thesecond array 64 could include between 43 and 56 stator vanes, and so on.

This could go on for as many interactions as required. In some casesonly a portion of the stages within a particular engine section would bedefined by this range. For instance, in a geared turbofan, in oneexample the first compressor stage may be cutoff while the second andthird stages would be modal, cuton, and follow the range defined above.It should be noted that the term “stage” as used in this disclosure mayrefer to a stator-rotor-stator combination.

The effect of selecting the relative number of airfoils using the factorF is illustrated schematically between FIGS. 2 and 3. With reference toFIG. 2, a modal, cuton arrangement is illustrated. In FIG. 2, the numberof airfoils 66 in the second array 64 is dictated by the factor F. Asthe rotor blades of the second array 64 rotate during operation of thegas turbine engine 20, a pulsing, unsteady pressure is distributed alongeach airfoil 66. The average direction of this pulsing, unsteadypressure is illustrated at U, and is known as the direction of unsteadylift. The average direction of the unsteady lift U is substantiallyperpendicular to the direction of the chord 68.

Because of the selected number of airfoils between the first array 60and the second array 64, noise (i.e., sound) propagates generally in thedirection N₁. As illustrated, the noise wave propagation direction N₁ issubstantially perpendicular to the direction of the unsteady lift U.Thus, the pulsating pressure waves of the unsteady lift U do not coupleeffectively with the noise waves. As a result, the amplitude of thenoise waves generated by this interaction is reduced, which in turnreduces noise.

FIG. 3 illustrates a non-modal, cuton arrangement. In FIG. 3, an examplelow pressure compressor 44′ includes a first, upstream array 60′ havinga number of airfoils that outnumbers those of the second, downstreamarray 64′. In this example, noise propagates in a direction N₂, which issubstantially parallel to the direction of unsteady lift U. In thisinstance, the pressure waves from the unsteady lift U couple better withthe noise waves than they do in the FIG. 2 arrangement, and thus thereis no significant noise reduction.

FIG. 4 graphically illustrates the relationship between sound powerlevel, in decibels (dB), and the factor F. As mentioned above, when“cutoff” is not possible, maximum noise reduction is achieved when thefactor F is within a range between 1.19 and 1.55. As shown above inEquation 1, the factor F is variable based on a number of factors. Oneparticular factor is airfoil shape (including, as examples,characteristics like airfoil camber and airfoil metal angle distributionalong the chord), which is captured by stagger angle θ in this example.FIG. 4 shows a plot of sound power level to the factor F for aparticular stagger angle θ. In this example, the largest amount of soundreduction is achieved at point 76, which corresponds to a factor F ofabout 1.25. This factor F is within the range between 1.19 and 1.55mentioned above.

As mentioned, this disclosure provides a modal, cuton airfoilarrangement, which achieves substantial noise reduction relative tonon-modal cuton arrangements. This disclosure has particular benefit insmall engines that cannot achieve a cutoff arrangement. As mentionedabove, this disclosure has benefits in other engines where certainstages can be “cutoff,” while others may benefit from a modal, cutoninteraction.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

One of ordinary skill in this art would understand that theabove-described embodiments are exemplary and non-limiting. That is,modifications of this disclosure would come within the scope of theclaims. Accordingly, the following claims should be studied to determinetheir true scope and content.

What is claimed is:
 1. A section of a gas turbine engine, comprising: afirst array of airfoils, the first array including a first number ofairfoils; and a second array of airfoils downstream of the first arrayof airfoils, the second array including a second number of airfoils,wherein the second number of airfoils is at least 1.19 times the firstnumber of airfoils thereby providing a predetermined modal.
 2. Thesection as recited in claim 1, wherein the second number of airfoils iswithin a range between 1.19 and 1.55 times the first number of airfoils.3. The section as recited in claim 2, wherein the range is defined bythe following equation ${1 + \frac{M_{tip}\sin \; \theta}{1 - M}};$and wherein M_(tip) is a tip rotational Mach number, M is the Machnumber into the second array from a frame of reference of the secondarray, and θ is a stagger angle of the second array.
 4. The section asrecited in claim 3, wherein the stagger angle is the incline of a chordbetween a leading edge and a trailing edge of an airfoil relative to adirection parallel to an engine central longitudinal axis.
 5. Thesection as recited in claim 3, wherein the range is further defined by:$\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$when the upstream array is a rotor and the downstream array is a stator;$\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$when the upstream array is a stator and the downstream array is a rotor;wherein n is the harmonic of blade passing frequency, and k is theharmonic of vane passing frequency.
 6. The section as recited in claim5, wherein n and k are equal to
 1. 7. The section as recited in claim 1,wherein the section includes a plurality of arrays of airfoils, andwherein a number of airfoils in each array is at least 1.19 times anumber of airfoils in an immediately upstream array.
 8. The section asrecited in claim 7, wherein the number of airfoils in each array iswithin a range between 1.19 and 1.55 times the number of airfoils in theimmediately upstream array.
 9. The section as recited in claim 1,wherein the first array of airfoils is an array of stator vanes, andwherein the second array of airfoils is an array of rotor blades. 10.The section as recited in claim 1, wherein the section is a low pressurecompressor.
 11. A gas turbine engine, comprising: a low pressurecompressor including a first array of airfoils and a second array ofairfoils downstream of the first array, the first array including afirst number of airfoils and the second array including a second numberof airfoils, wherein the second number of airfoils is at least 1.19times the first number of airfoils thereby providing a predeterminedmodal.
 12. The engine as recited in claim 11, wherein the second numberof airfoils is within a range between 1.19 and 1.55 times the firstnumber of airfoils.
 13. The engine as recited in claim 12, wherein therange is defined by the following equation${1 + \frac{M_{tip}\sin \; \theta}{1 - M}};$ and wherein M_(tip) is atip rotational Mach number, M is the Mach number into the second arrayfrom a frame of reference of the second array, and θ is a stagger angleof the second array.
 14. The engine as recited in claim 13, wherein thestagger angle is the incline of a chord between a leading edge and atrailing edge of an airfoil relative to a direction parallel to anengine central longitudinal axis.
 15. The engine as recited in claim 14,wherein the range is further defined by:$\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$when the upstream array is a rotor and the downstream array is a stator;$\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$when the upstream array is a stator and the downstream array is a rotor;wherein n is the harmonic of blade passing frequency, and k is theharmonic of vane passing frequency.
 16. The engine as recited in claim15, wherein n and k are equal to
 1. 17. A method, comprising:controlling an interaction between a first array of airfoils and asecond arrays of airfoils during operation of a gas turbine engine byproviding a number of airfoils in the second array that is at least 1.19times the number of airfoils in the first array, wherein the secondarray of airfoils is downstream of the first array of airfoils.
 18. Themethod as recited in claim 17, wherein the number of airfoils in thesecond array is within a range between 1.19 and 1.55 times the number ofairfoils in the first array.
 19. The method as recited in claim 18,wherein the range is defined by the following equation${1 + \frac{M_{tip}\sin \; \theta}{1 - M}};$ and wherein M_(tip) is atip rotational Mach number, M is the Mach number into the second arrayfrom a frame of reference of the second array, and θ is a stagger angleof the second array.
 20. The method as recited in claim 19, wherein therange is further defined by:$\frac{n}{k}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$when the upstream array is a rotor and the downstream array is a stator;$\frac{k}{n}( {1 + \frac{M_{tip}\sin \; \theta}{1 - M}} )$when the upstream array is a stator and the downstream array is a rotor;wherein n is the harmonic of blade passing frequency, and k is theharmonic of vane passing frequency.